Helicopter blade airfoil

ABSTRACT

A family of airfoil cross sections, termed SC21xx, for use in a helicopter blade is disclosed. The airfoil (36) achieves maximum lift performance equivalent to prior art airfoil configurations without incurring increased aerodynamic drag at high velocities. The airfoil (36) was developed by thickening and drooping the leading edge region (38) of the prior art airfoil (30) improving lift in the leading edge region (38) and delaying the formation of sonic shock waves at high velocities.

.Iadd.This application is a reissue of 06/903,169, U.S. Pat. No.4,744,728, now surrendered. .Iaddend.

FIELD OF THE INVENTION

The present invention relates to a blade for a helicopter, and moreparticularly, to the airfoil cross section thereof.

BACKGROUND

Rotorcraft, such as helicopters, are supported vertically by a pluralityof driven blades which rotate about a vertical axis. It will beappreciated that, since both the lift and propulsive force of thehelicopter are supplied primarily through the large rotating bladesystem, it is advantageous to provide a main blade configuration whichachieves a high lifting force at a given airspeed and which does notexperience a high aerodynamic drag.

The prior art contains many examples of blades and airfoils attemptingto achieve this high lift-low drag performance ideal, most notably U.S.Pat. No. 3,728,045 issued Apr. 17, 1973 to Balch, U.S. Pat. No.4,142,837 issued Mar. 6, 1979 to de Simone and U.S. Pat. No. 4,569,633issued Feb. 11, 1986 to Flemming, Jr.

The Balch airfoil, designated SC1095, provides an airfoil cross sectionwhich achieves both higher maximum lift and lower zero lift aerodynamicdrag at high velocity conditions as compared to the then-existingairfoils. These improvements were realized by selectively shaping theairfoil surface for both higher lift and lower drag by delayingseparation of the airflow boundary layer over the airfoil, as well asdecreasing the local Mach number at high freestream velocities.

The de Simone airfoil, also referred to as the SC1095-R8 configuration,is characterized in the referenced patent as an improvement over theSC1095 airfoil wherein the airfoil upper surface is shaped to distributethe surface static pressure peak over a greater area thereby reducingthe likelihood of airfoil boundary layer separation. The SC1095-R8airfoil achieves higher maximum lift than the SC1095 configuration atlower velocities, but is subject to higher zero lift drag forces at highvelocity conditions.

The Flemming, Jr. configuration is an improvement over the SC1095-R8design and is particularly well adapted for use in high speedapplications, such as in the radially outer tip portion of the mainblade. The Flemming, Jr. airfoil, also termed the SSC-Axx family,further reduces the zero lift drag force at high air speeds by delayingthe creation of shock waves in the local airflow. This further reductionin aerodynamic drag at high air speed is achieved at the expense of somemaximum lift at lower velocity operation.

As will be apparent from considering the above mentioned references as agroup, prior art blade designers combine several types of airfoil crosssections along the span of an individual blade in an attempt to maximizethe lift and minimize the drag over the range of expected airflowvelocities. For example, U.S. Pat. No. 4,248,572 issued Feb. 3, 1981 toFradenburgh shows a single helicopter blade which utilizes the high liftSC1095-R8 airfoil configuration in the lower velocity central spanregion of the blade and the lower lift SC1095 airfoil section radiallyoutward thereof in that portion of the blade which encounters higher airvelocities. The Flemming, Jr. patent provides a still furthermodification wherein members of the SSC-Axx family of airfoils is usedin the rotor tip portion due to its still greater resistance to shockwave formation at high airflow velocities.

As can be seen, the twin goals of high lift and low drag in prior artairfoils are exclusionary, leading designers to specify high liftairfoil configurations only in the lower speed regions of the rotorcraftblade while being content with reduced lift in the radially outer highspeed portions in order to avoid excessive aerodynamic drag. What isneeded is an airfoil configuration able to produce high lift withoutexperiencing unacceptably high drag force under high airflow velocityoperation.

SUMMARY OF THE INVENTION

It is therefore an object of the present invention to provided a bladefor a rotorcraft which exhibits improved performance in the form ofreduced aerodynamic drag while maintaining the maximum lift available inprior art blades.

It is further an object of the present invention to provide an improvedrotor airfoil cross section configured to delay the onset of aerodynamicdrag divergence as the airfoil is subjected to increasing velocityairflow over the blade exterior.

It is further an object of the present invention to provide an improvedrotor airfoil cross section which experiences a reduced quarter-chordpitching moment as compared to prior art high lift, low dragconfigurations, thereby reducing torsional stress within the blade.

It is further an object of the present invention to provide an improvedrotor airfoil cross section which is suited for use in the region of thehelicopter blade span extending from the rotor hub to a pointapproximately 90% of the distance to the blade tip.

It is still further an object of the present invention to provide afamily of similar shaped airfoil cross sections of differing thicknessesfor use along the blade span.

The present invention provides an improved airfoil cross section for usein a rotating helicopter blade or the like and which experiences reducedzero lift aerodynamic drag as compared to prior art airfoilconfigurations without incurring a penalty in the form of reducedmaximum lift. Improved performance is accomplished by altering theleading edge and surface geometry of a prior art, low drag airfoilconfiguration to increase the maximum lift coefficient withoutsignificantly changing the drag coefficient or pitching moment about thequarter chord point.

More specifically, the forward portion of a prior art SSC-Axx familyblade airfoil, the SSC-A10, is thickened and "drooped" slightly,improving the blade lift coefficient. The result is a high lift, lowdrag airfoil design which obtains the best advantages of prior artdesigns with a single configuration. The performance of the airfoilsection according to the present invention over the range of operatingconditions is such that the same family of airfoil shapes may be usedalong nearly the entire helicopter rotor blade span, preferably in thatportion of the span extending from the rotor hub to 90% of the spandistance between the hub and the blade tip. A rotorcraft equipped with arotor blade having an airfoil cross section according to the presentinvention requires less power to drive the main rotor, especially duringperiods of high speed horizontal flight in which the advancing rotorblades are subject to high relative velocity airflows.

Additionally, by providing a reduced quarter-chord pitching moment ascompared to prior art high lift blade airfoils, blades utilizingairfoils according to the present invention are subject to lesstorsional stress along the blade span thereby exhibiting less twistingbetween the hub and tip sections. Both these and other advantages andfeatures will be apparent to those skilled in the art upon close reviewof the following specification and the appended claims and drawingfigures.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 shows a schematic view of a rotating blade for a rotorcraft orthe like.

FIG. 2 shows a cross section of the airfoil according to the presentinvention and a prior art SSC-A10 airfoil cross section.

FIG. 3 is a graphical representation of the relationship between thecoefficient of zero lift aerodynamic drag and the Mach number forairfoils according to the present invention and a prior art SC1095-R8airfoil.

FIG. 4 is a graphical representation of the relationship between themaximum aerodynamic lift coefficient and Mach number for the airfoilaccording to the present invention and a prior art SC1095-R8 airfoil.

FIG. 5 is a graphical comparison of the maximum lift coefficient anddrag divergence number for an airfoil according to the present inventionand a series of prior art airfoils.

BEST MODE FOR CARRYING OUT THE INVENTION

FIG. 1 shows a schematic of a helicopter rotor 10 which consists of aplurality of equally spaced blades 12 mounted for rotation with a rotorhub 14 about an axis of rotation 16. Each blade 12 is preferablyidentical to the single blade illustrated in FIG. 1 and includes a rootportion 18 which connects to the hub 14, a tip portion 20 which is thatpart of the blade farthest from the axis of rotation 16 and whichtherefore travels at the highest speed, and a center portion 22extending between the hub 14 and tip 20. Each of the blades 12 has aleading edge 24, a trailing edge 26 and defines a chord dimension C anda span dimension S as shown in FIG. 1. The blade 12 is an airfoil incross section and generates lift during rotation of the blade 12.

Turning now to FIG. 2, a cross section 28 of the airfoil according tothe present invention is shown overlain with a cross section 30 of aprior art SSC-A10 airfoil. The airfoil according to the presentinvention, also termed "SC2110", is shown with the leading edge 24 andtrailing edge 26 defining a chordal distance C therebetween. FIG. 2 alsoshows the airfoil chord line 30 connecting the leading and trailingedges 24, 26. The airfoil section 28 includes an upper surface 34 and alower surface 36 disposed on opposite sides of the chord line 32.

It is conventional to define an individual airfoil shape by settingforth the location of the upper airfoil surface 34 and the lower airfoilsurface 36 at a plurality of stations disposed along the blade chord 32.The SC2110 airfoil cross section 28 according to the present inventionis defined in Table I which contains a listing of the verticaldisplacement Y_(u), Y_(l) of the airfoil upper and lower surfaces 34, 36as a fraction of the airfoil chordal dimension C. The horizontaldisplacement of each station is represented by the column titled X/C andrepresents the proportional displacement along the airfoil chord line 32from the leading edge 24 to the trailing edge 26.

                  TABLE I                                                         ______________________________________                                        X/C            Y.sub.u/c                                                                              Y.sub.l/c                                             ______________________________________                                        0.0            0.0      0.0                                                   0.004179       0.012216 -0.007622                                             0.006589       0.015843 -0.008855                                             0.011469       0.021681 -0.010655                                             0.023818       0.032107 -0.013265                                             0.048677       0.044925 -0.016700                                             0.073637       0.052653 -0.019156                                             0.098647       0.057671 -0.021112                                             0.148716       0.063313 -0.024060                                             0.198856       0.065816 -0.026699                                             0.249026       0.066705 -0.029199                                             0.299206       0.066734 -0.031285                                             0.337529       0.066424 -0.032514                                             0.377406       0.065877 -0.033372                                             0.417281       0.065058 -0.033801                                             0.437217       0.064511 -0.033848                                             0.457152       0.063848 -0.033784                                             0.477085       0.063055 -0.033605                                             0.497018       0.062135 -0.033304                                             0.536880       0.059927 -0.032343                                             0.556810       0.058641 -0.031680                                             0.576739       0.057228 -0.030903                                             0.596668       0.055676 -0.030019                                             0.636519       0.052112 -0.027955                                             0.656445       0.050076 -0.026779                                             0.676366       0.047861 -0.025512                                             0.696288       0.045461 -0.024157                                             0.736125       0.040116 -0.021213                                             0.756042       0.037186 -0.011641                                             0.775956       0.034105 -0.018017                                             0.795871       0.030894 -0.016351                                             0.835698       0.024166 -0.012929                                             0.855611       0.020691 -0.011180                                             0.875522       0.017178 -0.009404                                             0.895435       0.013670 -0.007600                                             0.935263       0.007073 -0.004043                                             0.955180       0.004381 -0.002452                                             0.975106       0.002582 -0.001260                                             0.985074       0.002218 -0.000912                                             0.995045       0.002365 -0.000821                                             1.000000       0.002676 -0.000899                                             ______________________________________                                    

As with the SSC-A10 airfoil, the SC2110 airfoil is a member of acorresponding family of similar shapes of differing thicknesses asdefined by the ratio of airfoil maximum thickness (t_(max)) to airfoilchord (C). The airfoil designations SSC-A10 and SC2110 thus eachrepresent individual members of the respective airfoil families SSC-Axxand SC21xx wherein the ratio of t_(max) to C is approximately 0.10, or10%. The SC21xx family of airfoils is defined by Table II wherein thevertical displacement of the upper and lower airfoil surfaces istabulated at a plurality of stations along the airfoil chord 32, eachstation defined as in Table I as a proportional displacement X/C of thedistance between the blade leading and trailing edges 24, 26. Verticaldisplacements are expressed as a ratio of the vertical displacementY_(u), Y_(l), to the blade maximum thickness, t_(max).

                  TABLE II                                                        ______________________________________                                        X/C            Y.sub.u /t.sub.max                                                                     Y.sub.l /t.sub.max                                    ______________________________________                                        0.000000       0.000000  0.000000                                             0.004179       0.123145 -0.076835                                             0.006589       0.159708 -0.089264                                             0.011469       0.218559 -0.107409                                             0.023818       0.323659 -0.133720                                             0.048677       0.452873 -0.168347                                             0.073637       0.530776 -0.193105                                             0.098647       0.581361 -0.212823                                             0.148716       0.638236 -0.242540                                             0.198856       0.663468 -0.269143                                             0.249026       0.672430 -0.294345                                             0.299206       0.672722 -0.315373                                             0.337529       0.669597 -0.327762                                             0.377406       0.664083 -0.336411                                             0.417281       0.655827 -0.340736                                             0.437217       0.650313 -0.341210                                             0.457152       0.643629 -0.340565                                             0.477085       0.635635 -0.338760                                             0.497018       0.626361 -0.335726                                             0.536880       0.604103 -0.326038                                             0.556810       0.591139 -0.319355                                             0.576739       0.576895 -0.311522                                             0.596668       0.561250 -0.302611                                             0.636519       0.525323 -0.281805                                             0.656445       0.504799 -0.269950                                             0.676366       0.482470 -0.257178                                             0.696288       0.458276 -0.243518                                             0.736125       0.404395 -0.213841                                             0.756042       0.374859 -0.117349                                             0.775956       0.343801 -0.181623                                             0.795871       0.311432 -0.164829                                             0.835698       0.243609 -0.130333                                             0.855611       0.208579 -0.112702                                             0.875522       0.173165 -0.094798                                             0.895435       0.137802 -0.076613                                             0.935263       0.071300 -0.040756                                             0.955180       0.044163 -0.024718                                             0.975106       0.026028 -0.012702                                             0.985074       0.022359 -0.009194                                             0.995045       0.023841 -0.008276                                             1.000000       0.026976 -0.009063                                             ______________________________________                                    

It has been determined that the advantages of increased performance arestill achieved by the SC21xx airfoils generally, and the SC2110 airfoilparticularly, it the tabulated quantities vary throughout a rang of ±three percent.

The airfoil section 28 according to the present invention is aderivative of the SSC-A10 cross section 30 and differs therefrom both inconfiguration and in performance. The SC2110 cross section 28 issignificantly thicker than the prior art airfoil section 30 in theleading edge region 38 and the central chord region 40, while beingsubstantially similar in thickness in the trailing edge region 42. Theleading edge, central chord, and trailing edge regions 38, 40, 42, areas defined in U.S. Pat. No. 4,569,633 issued to Flemming, Jr., discussedin the Background section hereinabove.

It should further be noted that this increased thickness is accompaniedby an increased camber or "droop" in the forward portions 38, 40 of theairfoil cross section, resulting in the upper surface 34 of the SC2110cross section 28 being located at a greater vertical displacement fromthe airfoil chord line 32 as compared to the corresponding surface onthe prior art SSC-A10 cross section 30. This droop is most accuratelyexpressed as rotation of the prior art SSC-A10 nose of about 3.25° aboutthe X/C=0.15 chord station.

The drooped configuration of the present invention provides additionallift over the prior art SSC-A10 airfoil 30, restoring lift performancenearly to that of the SC1095-R8 configuration as discussed in theBackground section. The reshaped cross section also evens out thevariation of local air velocity and Mach number over the forward surfaceof the airfoil 30, thereby delaying the formation of sonic shock waves.

The success of the SC2110 airfoil section 28 is evident from FIGS. 3 and4 which compare the lift and drag coefficients of individual blades overa range of Mach numbers. FIG. 4 shows the coefficient of maximum lift,C_(lmax), for a prior art SC1095-R8 airfoil 44 to a similar curve 46representing the performance of a blade utilizing an airfoil accordingto the present invention. The two curves 44, 46 are nearly identicalover the range of Mach numbers 0.3 to 0.6 indicating that liftperformance of the SC2110 is comparable to the SC1095-R8 airfoil. TheSC1095-R8 airfoil has been selected for the comparison of FIGS. 3 and 4as it is the highest lift rotorcraft airfoil configuration known in theprior art.

FIG. 3 clearly shows the greatest single advantage of the airfoilaccording to the present invention over the prior art SC1095-R8 highlift airfoil. FIG. 3 is a plot of the drag coefficient of individualblades at the zero lift orientation, C_(do), over the range of Machnumbers 0.7 to 0.86. The curve of the prior art SC1095-R8 airfoil 48lies well above the curve 50 which represents the performance of theairfoil according to the present invention. By comparing FIGS. 3 and 4,it should be readily apparent that a blade utilizing an airfoilaccording to the present invention provides lifting performancesubstantially equivalent to that of the best rotorcraft airfoil sectionsheretofore known in the art while simultaneously inducing less drag inthe upper Mach number operating range.

More significantly, the airfoil of the present invention can operate ata significantly higher Mach number before reaching the drag divergenceair speed limit above which aerodynamic drag increases rapidly andunacceptably. The drag divergence is defined as the air speed at whichthe ratio of the change of the zero lift drag coefficient to the changeof Mach number reaches a value of 0.1. These points are shown in FIG. 3for the prior art airfoil 52 and the airfoil according to the presentinvention 54 illustrating the advantage of the SC2110 airfoil over theprior art high lift airfoils at high Mach number operation.

FIG. 5 presents a comparison of the lift and drag performance of theSC2110 airfoil with that of its predecessor configurations. FIG. 5 is aplot of each airfoil section wherein the vertical displacementrepresents the maximum lift coefficient at a relatively low Mach number,M=0.3, and the horizontal displacement is equivalent to the dragdivergence Mach number (M_(dd)). As will be appreciated by those skilledin the art, the point 56 representing the airfoil according to thepresent invention has pushed the operating envelope for rotorcraftblades significantly toward the upper right hand corner as compared tothe performance of the prior art SC1095, SC1095-R8 and SSC-Axx familyairfoils 58, 60, 62, signifying an airfoil section which will achievehigh lift force over nearly the entire blade airspeed operating range.

The broad opening range of the airfoil according to the presentinvention allows a blade 12 to utilize an individual airfoil sectionover nearly the entire span of the blade, unlike prior art bladeswherein airfoil sections had to be matched with expected Mach number inorder to avoid undue drag and undesirably high power consumption. Theairfoil according to the present invention is thus able to be used overthat portion of the blade span extending from the central hub 14 tonearly 90% of the span length toward the blade tip 20, providing thehigh maximum lift characteristics of the SC1095-R8 configuration withoutexperiencing the drag divergence phenomenon inherent in the prior artairfoil shapes.

The remaining outer 10% of the blade span may utilize a prior artSSC-Axx airfoil or other configuration best suited for the high speedtip region wherein the generation of lift is of secondary importance ascompared to actual drag.

It is thus apparent that the airfoil cross section 28 according to thepresent invention is well suited to achieve the objects and advantagesas set forth hereinabove, and it will further be appreciated by thoseskilled in the art that various minor modifications may be made from theconfiguration as presented without departing from the spirit and scopeof the invention as described and claimed.

We claim:
 1. A blade for a rotorcraft having a first root end mounted toa central hub, a second tip end, and a span extending between the rootand tip ends, the blade further including a high lift, low drag airfoilcross section having a leading edge and a trailing edge, an upperairfoil surface and a lower airfoil surface, each surface extendingbetween the leading and trailing edges and characterized by thetabulation:

    ______________________________________                                        X/C           Y.sub.u/c                                                                              Y.sub.l/e                                              ______________________________________                                        0.0           0.0      0.0                                                    0.004179      0.012216 -0.007622                                              0.006589      0.015843 -0.008855                                              0.011469      0.021681 -0.010655                                              0.023818      0.032107 -0.013265                                              0.048677      0.044925 -0.016700                                              0.073637      0.052653 -0.019156                                              0.098647      0.057671 -0.021112                                              0.148716      0.063313 -0.024060                                              0.198856      0.065816 -0.026699                                              0.249026      0.066705 -0.029199                                              0.299206      0.066734 -0.031285                                              0.337529      0.066424 -0.032514                                              0.377406      0.065877 -0.033372                                              0.417281      0.065058 -0.033801                                              0.437217      0.064511 -0.033848                                              0.457152      0.063848 -0.033784                                              0.477085      0.063055 -0.033605                                              0.497018      0.062135 -0.033304                                              0.536880      0.059927 -0.032343                                              0.556810      0.058641 -0.031680                                              0.576739      0.057228 -0.030903                                              0.596668      0.055676 -0.030019                                              0.636519      0.052112 -0.027955                                              0.656445      0.050076 -0.026779                                              0.676366      0.047861 -0.025512                                              0.696288      0.045461 -0.024157                                              0.736125      0.040116 -0.021213                                              0.756042      0.037186 -0.011641                                              0.775956      0.034105 -0.018017                                              0.795871      0.030894 -0.016351                                              0.835698      0.024166 -0.012929                                              0.855611      0.020691 -0.011180                                              0.875522      0.017178 -0.009404                                              0.895435      0.013670 -0.007600                                              0.935263      0.007073 -0.004043                                              0.955180      0.004381 -0.002462                                              0.975106      0.002582 -0.001260                                              0.985074      0.002218 -0.000912                                              0.995045      0.002365 -0.000821                                              1.000000      0.002676 -0.000899                                              ______________________________________                                    

wherein X is the linear displacement along a chord line extendingbetween the airfoil leading edge and the airfoil trailing edge; C is thechordal length of the airfoil cross section measured between the leadingand the trailing edges; Y_(u) is the transverse displacement of theairfoil upper surface from the chord line; and Y_(l) is the transversedisplacement of the airfoil lower surface from the chord line.
 2. Therotorcraft blade as recited in claim 1, whereinthe high lift-low dragairfoil section is defined within that portion of the blade extendingspanwisely from the root end to at least 90% of the distance to the tipend.
 3. The rotorcraft blade as recited in claim 1, wherein the valuesof Y_(l/C) and Y_(u/C) are within 3% of the tabulated values.
 4. Afamily of airfoil cross sections for a helicopter blade each having aleading edge and a trailing edge, an upper airfoil surface and a lowerairfoil surface, each surface extending between the leading and trailingedges and characterized by the tabulation:

    ______________________________________                                        X/C            Y.sub.u /t.sub.max                                                                     Y.sub.l /t.sub.max                                    ______________________________________                                        0.000000       0.000000  0.000000                                             0.004179       0.123145 -0.076835                                             0.006589       0.159708 -0.089264                                             0.011469       0.218559 -0.107409                                             0.023818       0.323659 -0.133720                                             0.048677       0.452873 -0.168347                                             0.073637       0.530776 -0.193105                                             0.098647       0.581361 -0.212823                                             0.148716       0.638236 -0.242540                                             0.198856       0.663468 -0.269143                                             0.249026       0.672430 -0.294345                                             0.299206       0.672722 -0.315373                                             0.337529       0.669597 -0.327762                                             0.377406       0.664083 -0.336411                                             0.417281       0.655827 -0.340736                                             0.437217       0.650313 -0.341210                                             0.457152       0.643629 -0.340565                                             0.477085       0.635635 -0.338760                                             0.497018       0.626361 -0.335726                                             0.536880       0.604103 -0.326038                                             0.556810       0.591139 -0.319355                                             0.576739       0.576895 -0.311522                                             0.596668       0.561250 -0.302611                                             0.636519       0.525323 -0.281805                                             0.656445       0.504799 -0.269950                                             0.676366       0.482470 -0.257178                                             0.696288       0.458276 -0.243518                                             0.736125       0.404395 -0.213841                                             0.756042       0.374859 -0.117349                                             0.775956       0.343801 -0.181623                                             0.795871       0.311432 -0.164829                                             0.835698       0.243609 -0.130333                                             0.855611       0.208579 -0.112702                                             0.875522       0.173165 -0.094798                                             0.895435       0.137802 -0.076613                                             0.935263       0.071300 -0.040756                                             0.955180       0.044163 -0.024718                                             0.975106       0.026028 -0.012702                                             0.985074       0.022359 -0.009194                                             0.995045       0.023841 -0.008276                                             1.000000       0.026976 -0.009063                                             ______________________________________                                    

wherein X is the linear displacement along a chord line extendingbetween the airfoil leading edge and the airfoil trailing edge; C is thechordal length of the airfoil cross section measured between the leadingand trailing edges; Y_(u) is the transverse displacement of the airfoilupper surface from the chord line; Y_(l) is the transverse displacementof the airfoil lower surface from the chord line; and t_(max) is themaximum transverse thickness achieved by the airfoil cross section. 5.The family of airfoil cross sections as recited in claim 4, wherein thevalues of Y_(u) /t_(max) and Y_(l) /t_(max) are within 3% of thetabulated values.